Geared turbofan gas turbine engine architecture

ABSTRACT

A gas turbine engine includes a very high speed low pressure turbine such that a quantity defined by the exit area of the low pressure turbine multiplied by the square of the low pressure turbine rotational speed compared to the same parameters for the high pressure turbine is at a ratio between about 0.5 and about 1.5.

CROSS REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. patent application Ser. No.14/662,387, filed Mar. 19, 2015, which is a continuation-in-part of U.S.application Ser. No. 13/629,681 filed on Sep. 28, 2012, which is acontinuation-in-part of U.S. application Ser. No. 13/363,154 filed onJan. 31, 2012.

BACKGROUND

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section and a turbine section. Air entering thecompressor section is compressed and delivered into the combustionsection where it is mixed with fuel and ignited to generate a high-speedexhaust gas flow. The high-speed exhaust gas flow expands through theturbine section to drive the compressor and the fan section. Thecompressor section typically includes low and high pressure compressors,and the turbine section includes low and high pressure turbines.

The high pressure turbine drives the high pressure compressor through anouter shaft to form a high spool, and the low pressure turbine drivesthe low pressure compressor through an inner shaft to form a low spool.The inner shaft may also drive the fan section. A direct drive gasturbine engine includes a fan section driven by the inner shaft suchthat the low pressure compressor, low pressure turbine and fan sectionrotate at a common speed in a common direction.

A speed reduction device such as an epicyclical gear assembly may beutilized to drive the fan section such that the fan section may rotateat a speed different than the turbine section so as to increase theoverall propulsive efficiency of the engine. In such enginearchitectures, a shaft driven by one of the turbine sections provides aninput to the epicyclical gear assembly that drives the fan section at aspeed different than the turbine section such that both the turbinesection and the fan section can rotate at closer to optimal speeds.

Although geared architectures have improved propulsive efficiency,turbine engine manufacturers continue to seek further improvements toengine performance including improvements to thermal, transfer andpropulsive efficiencies.

SUMMARY

In a featured embodiment, a turbine section of a gas turbine engine hasa first turbine section, and a second turbine section, wherein the firstturbine section has a first exit area at a first exit point and rotatesat a first speed. The second turbine section has a second exit area at asecond exit point and rotates at a second speed, which is faster thanthe first speed. A first performance quantity is defined as the productof the first speed squared and the first area. A second performancequantity is defined as the product of the second speed squared and thesecond area. A ratio of the first performance quantity to the secondperformance quantity is between about 0.5 and about 1.5.

In another embodiment according to the previous embodiment, the ratio isabove or equal to about 0.8.

In another embodiment according to the previous embodiment, the firstturbine section has at least 3 stages.

In another embodiment according to the previous embodiment, the firstturbine section has up to 6 stages.

In another embodiment according to the previous embodiment, the secondturbine section has 2 or fewer stages.

In another embodiment according to the previous embodiment, a pressureratio across the first turbine section is greater than about 5:1.

In another featured embodiment, a gas turbine engine has a fan, acompressor section in fluid communication with the fan, a combustionsection in fluid communication with the compressor section, and aturbine section in fluid communication with the combustion section. Theturbine section includes a first turbine section and a second turbinesection. The first turbine section has a first exit area at a first exitpoint and rotates at a first speed. The second turbine section has asecond exit area at a second exit point and rotates at a second speed,which is higher than the first speed. A first performance quantity isdefined as the product of the first speed squared and the first area. Asecond performance quantity is defined as the product of the secondspeed squared and the second area. A ratio of the first performancequantity to the second performance quantity is between about 0.5 andabout 1.5.

In another embodiment according to the previous embodiment, the ratio isabove or equal to about 0.8.

In another embodiment according to the previous embodiment, thecompressor section includes a first compressor section and a secondcompressor section, wherein the first turbine section and the firstcompressor section rotate in a first direction, and wherein the secondturbine section and the second compressor section rotate in a secondopposed direction.

In another embodiment according to the previous embodiment, a gearreduction is included between the fan and a low spool driven by thefirst turbine section such that the fan rotates at a lower speed thanthe first turbine section.

In another embodiment according to the previous embodiment, the fanrotates in the second opposed direction.

In another embodiment according to the previous embodiment, the gearreduction is greater than about 2.3.

In another embodiment according to the previous embodiment, the gearratio is greater than about 2.5.

In another embodiment according to the previous embodiment, the ratio isabove or equal to about 1.0.

In another embodiment according to the previous embodiment, the fandelivers a portion of air into a bypass duct, and a bypass ratio beingdefined as the portion of air delivered into the bypass duct divided bythe amount of air delivered into the first compressor section, with thebypass ratio being greater than about 6.0.

In another embodiment according to the previous embodiment, the bypassratio is greater than about 10.0.

In another embodiment according to the previous embodiment, the fan has26 or fewer blades.

In another embodiment according to the previous embodiment, the firstturbine section has at least 3 stages.

In another embodiment according to the previous embodiment, the firstturbine section has up to 6 stages.

In another embodiment according to the previous embodiment, a pressureratio across the first turbine section is greater than about 5:1.

Although the different examples have the specific components shown inthe illustrations, embodiments of this disclosure are not limited tothose particular combinations. It is possible to use some of thecomponents or features from one of the examples in combination withfeatures or components from another one of the examples.

These and other features disclosed herein can be best understood fromthe following specification and drawings, the following of which is abrief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a schematic view of an example gas turbine engine.

FIG. 2 is a schematic view indicating relative rotation between sectionsof an example gas turbine engine.

FIG. 3 is another schematic view indicating relative rotation betweensections of an example gas turbine engine.

FIG. 4 is another schematic view indicating relative rotation betweensections of an example gas turbine engine.

FIG. 5 is another a schematic view indicating relative rotation betweensections of an example gas turbine engine.

FIG. 6 is a schematic view of a bearing configuration supportingrotation of example high and low spools of the example gas turbineengine.

FIG. 7 is another schematic view of a bearing configuration supportingrotation of example high and low spools of the example gas turbineengine.

FIG. 8A is another schematic view of a bearing configuration supportingrotation of example high and low spools of the example gas turbineengine.

FIG. 8B is an enlarged view of the example bearing configuration shownin FIG. 8A.

FIG. 9 is another schematic view of a bearing configuration supportingrotation of example high and low spools of the example gas turbineengine.

FIG. 10 is a schematic view of an example compact turbine section.

FIG. 11 is a schematic cross-section of example stages for the disclosedexample gas turbine engine.

FIG. 12 is a schematic view an example turbine rotor perpendicular tothe axis or rotation.

FIG. 13 is another embodiment of an example gas turbine engine for usewith the present invention.

FIG. 14 is yet another embodiment of an example gas turbine engine foruse with the present invention.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates an example gas turbine engine 20 thatincludes a fan section 22, a compressor section 24, a combustor section26 and a turbine section 28. Alternative engines might include anaugmenter section (not shown) among other systems or features. The fansection 22 drives air along a bypass flow path B while the compressorsection 24 draws air in along a core flow path C where air is compressedand communicated to a combustor section 26. In the combustor section 26,air is mixed with fuel and ignited to generate a high pressure exhaustgas stream that expands through the turbine section 28 where energy isextracted and utilized to drive the fan section 22 and the compressorsection 24.

Although the disclosed non-limiting embodiment depicts a turbofan gasturbine engine, it should be understood that the concepts describedherein are not limited to use with turbofans as the teachings may beapplied to other types of turbine engines; for example a turbine engineincluding a three-spool architecture in which three spoolsconcentrically rotate about a common axis such that a low spool enablesa low pressure turbine to drive a fan via a gearbox, an intermediatespool enables an intermediate pressure turbine to drive a firstcompressor of the compressor section, and a high spool enables a highpressure turbine to drive a high pressure compressor of the compressorsection.

The example engine 20 generally includes a low speed spool 30 and a highspeed spool 32 mounted for rotation about an engine central longitudinalaxis A relative to an engine static structure 36 via several bearingsystems 38. It should be understood that various bearing systems 38 atvarious locations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatconnects a fan 42 and a low pressure (or first) compressor section 44 toa low pressure (or first) turbine section 46. The inner shaft 40 drivesthe fan 42 through a speed change device, such as a geared architecture48, to drive the fan 42 at a lower speed than the low speed spool 30.The high-speed spool 32 includes an outer shaft 50 that interconnects ahigh pressure (or second) compressor section 52 and a high pressure (orsecond) turbine section 54. The inner shaft 40 and the outer shaft 50are concentric and rotate via the bearing systems 38 about the enginecentral longitudinal axis A.

A combustor 56 is arranged between the high pressure compressor 52 andthe high pressure turbine 54. In one example, the high pressure turbine54 includes at least two stages to provide a double stage high pressureturbine 54. In another example, the high pressure turbine 54 includesonly a single stage. As used herein, a “high pressure” compressor orturbine experiences a higher pressure than a corresponding “lowpressure” compressor or turbine.

The example low pressure turbine 46 has a pressure ratio that is greaterthan about 5. The pressure ratio of the example low pressure turbine 46is measured prior to an inlet of the low pressure turbine 46 as relatedto the pressure measured at the outlet of the low pressure turbine 46prior to an exhaust nozzle.

A mid-turbine frame 58 of the engine static structure 36 is arrangedgenerally between the high pressure turbine 54 and the low pressureturbine 46. The mid-turbine frame 58 further supports bearing systems 38in the turbine section 28 as well as setting airflow entering the lowpressure turbine 46.

The core airflow C is compressed by the low pressure compressor 44 thenby the high pressure compressor 52 mixed with fuel and ignited in thecombustor 56 to produce high speed exhaust gases that are then expandedthrough the high pressure turbine 54 and low pressure turbine 46. Themid-turbine frame 58 includes vanes 60, which are in the core airflowpath and function as an inlet guide vane for the low pressure turbine46. Utilizing the vane 60 of the mid-turbine frame 58 as the inlet guidevane for low pressure turbine 46 decreases the length of the lowpressure turbine 46 without increasing the axial length of themid-turbine frame 58. Reducing or eliminating the number of vanes in thelow pressure turbine 46 shortens the axial length of the turbine section28. Thus, the compactness of the gas turbine engine 20 is increased anda higher power density may be achieved.

The disclosed gas turbine engine 20 in one example is a high-bypassgeared aircraft engine. In a further example, the gas turbine engine 20includes a bypass ratio greater than about six (6), with an exampleembodiment being greater than about ten (10). The example gearedarchitecture 48 is an epicyclical gear train, such as a planetary gearsystem, star gear system or other known gear system, with a gearreduction ratio of greater than about 2.3.

In one disclosed embodiment, the gas turbine engine 20 includes a bypassratio greater than about ten (10:1) and the fan diameter issignificantly larger than an outer diameter of the low pressurecompressor 44. It should be understood, however, that the aboveparameters are only exemplary of one embodiment of a gas turbine engineincluding a geared architecture and that the present disclosure isapplicable to other gas turbine engines.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft., withthe engine at its best cruise fuel consumption relative to the thrust itproduces—also known as “bucket cruise Thrust Specific Fuel Consumption(‘TSFC’)”—is the industry standard parameter of pound-mass (lbm) of fuelper hour being burned divided by pound-force (lbf) of thrust the engineproduces at that minimum bucket cruise point.

“Low fan pressure ratio” is the pressure ratio across the fan bladealone, without a Fan Exit Guide Vane (“FEGV”) system. The low fanpressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.50. In another non-limiting embodimentthe low fan pressure ratio is less than about 1.45.

“Low corrected fan tip speed” is the actual fan tip speed in ft/secdivided by an industry standard temperature correction of [(Tram °R)/(518.7°R)]^(0.5). The “Low corrected fan tip speed”, as disclosedherein according to one non-limiting embodiment, is less than about 1150ft/second.

The example gas turbine engine includes the fan 42 that comprises in onenon-limiting embodiment less than about 26 fan blades. In anothernon-limiting embodiment, the fan section 22 includes less than about 18fan blades. Moreover, in one disclosed embodiment the low pressureturbine 46 includes no more than about 6 turbine stages schematicallyindicated at 34. In another non-limiting example embodiment the lowpressure turbine 46 includes about 3 or more turbine stages. A ratiobetween the number of fan blades 42 and the number of low pressureturbine stages is between about 2.5 and about 8.5. The example lowpressure turbine 46 provides the driving power to rotate the fan section22 and therefore the relationship between the number of turbine stages34 in the low pressure turbine 46 and the number of blades 42 in the fansection 22 disclose an example gas turbine engine 20 with increasedpower transfer efficiency.

Increased power transfer efficiency is provided due in part to theincreased use of improved turbine blade materials and manufacturingmethods such as directionally solidified castings, and single crystalmaterials that enable increased turbine speed and a reduced number ofstages. Moreover, the example low pressure turbine 46 includes improvedturbine disks configurations that further enable desired durability atthe higher turbine speeds.

Referring to FIGS. 2 and 3, an example disclosed speed change device isan epicyclical gearbox of a planet type, where the input is to thecenter “sun” gear 62. Planet gears 64 (only one shown) around the sungear 62 rotate and are spaced apart by a carrier 68 that rotates in adirection common to the sun gear 62. A ring gear 66, which isnon-rotatably fixed to the engine static casing 36 (shown in FIG. 1),contains the entire gear assembly. The fan 42 is attached to and drivenby the carrier 68 such that the direction of rotation of the fan 42 isthe same as the direction of rotation of the carrier 68 that, in turn,is the same as the direction of rotation of the input sun gear 62.

In the following figures nomenclature is utilized to define the relativerotations between the various sections of the gas turbine engine 20. Thefan section is shown with a “+” sign indicating rotation in a firstdirection. Rotations relative to the fan section 22 of other features ofthe gas turbine engine are further indicated by the use of either a “+”sign or a “−” sign. The “−” sign indicates a rotation that is counter tothat of any component indicated with a “+” sign.

Moreover, the term fan drive turbine is utilized to indicate the turbinethat provides the driving power for rotating the blades 42 of the fansection 22. Further, the term “second turbine” is utilized to indicatethe turbine before the fan drive turbine that is not utilized to drivethe fan 42. In this disclosed example, the fan drive turbine is the lowpressure turbine 46, and the second turbine is the high pressure turbine54. However, it should be understood that other turbine sectionconfigurations that include more than the shown high and low pressureturbines 54, 46 are within the contemplation of this disclosure. Forexample, a three spool engine configuration may include an intermediateturbine (not shown) utilized to drive the fan section 22 and is withinthe contemplation of this disclosure.

In one disclosed example embodiment (FIG. 2) the fan drive turbine isthe low pressure turbine 46 and therefore the fan section 22 and lowpressure turbine 46 rotate in a common direction as indicated by thecommon “+” sign indicating rotation of both the fan 42 and the lowpressure turbine 46. Moreover in this example, the high pressure turbine54 or second turbine rotates in a direction common with the fan driveturbine 46. In another example shown in FIG. 3, the high pressureturbine 54 or second turbine rotates in a direction opposite the fandrive turbine (low pressure turbine 46) and the fan 42.

Counter rotating the low pressure compressor 44 and the low pressureturbine 46 relative to the high pressure compressor 52 and the highpressure turbine 54 provides certain efficient aerodynamic conditions inthe turbine section 28 as the generated high speed exhaust gas flowmoves from the high pressure turbine 54 to the low pressure turbine 46.The relative rotations in the compressor and turbine sections provideapproximately the desired airflow angles between the sections, whichimproves overall efficiency in the turbine section 28, and provides areduction in overall weight of the turbine section 28 by reducing oreliminating airfoils or an entire row of vanes.

Referring to FIGS. 4 and 5, another example disclosed speed changedevice is an epicyclical gearbox referred to as a star type gearbox,where the input is to the center “sun” gear 62. Star gears 65 (only oneshown) around the sun gear 62 rotate in a fixed position around the sungear and are spaced apart by a carrier 68 that is fixed to a staticcasing 36 (best shown in FIG. 1). A ring gear 66 that is free to rotatecontains the entire gear assembly. The fan 42 is attached to and drivenby the ring gear 66 such that the direction of rotation of the fan 42 isopposite the direction of rotation of the input sun gear 62.Accordingly, the low pressure compressor 44 and the low pressure turbine46 rotate in a direction opposite rotation of the fan 42.

In one disclosed example embodiment shown in FIG. 4, the fan driveturbine is the low pressure turbine 46 and therefore the fan 42 rotatesin a direction opposite that of the low pressure turbine 46 and the lowpressure compressor 44. Moreover in this example the high spool 32including the high pressure turbine 54 and the high pressure compressor52 rotate in a direction counter to the fan 42 and common with the lowspool 30 including the low pressure compressor 44 and the fan driveturbine 46.

In another example gas turbine engine shown in FIG. 5, the high pressureor second turbine 54 rotates in a direction common with the fan 42 andcounter to the low spool 30 including the low pressure compressor 44 andthe fan drive turbine 46.

Referring to FIG. 6, the bearing assemblies near the forward end of theshafts in the engine at locations 70 and 72, which bearings supportrotation of the inner shaft 40 and the outer shaft 50, counter netthrust forces in a direction parallel to the axis A that are generatedby the rearward load of low pressure turbine 46 and the high pressureturbine 54, minus the high pressure compressor 52 and the low pressurecompressor 44, which also contribute to the thrust forces acting on thecorresponding low spool 30 and the high spool 32.

In this example embodiment, a first forward bearing assembly 70 issupported on a portion of the static structure schematically shown at 36and supports a forward end of the inner shaft 40. The example firstforward bearing assembly 70 is a thrust bearing and controls movement ofthe inner shaft 40 and thereby the low spool 30 in an axial direction. Asecond forward bearing assembly 72 is supported by the static structure36 to support rotation of the high spool 32 and substantially preventmovement along in an axial direction of the outer shaft 50. The firstforward bearing assembly 70 is mounted to support the inner shaft 40 ata point forward of a connection 88 of a low pressure compressor rotor90. The second forward bearing assembly 72 is mounted forward of aconnection referred to as a hub 92 between a high pressure compressorrotor 94 and the outer shaft 50. A first aft bearing assembly 74supports the aft portion of the inner shaft 40. The first aft bearingassembly 74 is a roller bearing and supports rotation, but does notprovide resistance to movement of the shaft 40 in the axial direction.Instead, the aft bearing 74 allows the shaft 40 to expand thermallybetween its location and the bearing 72. The example first aft bearingassembly 74 is disposed aft of a connection hub 80 between a lowpressure turbine rotor 78 and the inner shaft 40. A second aft bearingassembly 76 supports the aft portion of the outer shaft 50. The examplesecond aft bearing assembly 76 is a roller bearing and is supported by acorresponding static structure 36 through the mid turbine frame 58 whichtransfers the radial load of the shaft across the turbine flow path toground 36. The second aft bearing assembly 76 supports the outer shaft50 and thereby the high spool 32 at a point aft of a connection hub 84between a high pressure turbine rotor 82 and the outer shaft 50.

In this disclosed example, the first and second forward bearingassemblies 70, 72 and the first and second aft bearing assemblies 74, 76are supported to the outside of either the corresponding compressor orturbine connection hubs 80, 88 to provide a straddle supportconfiguration of the corresponding inner shaft 40 and outer shaft 50.The straddle support of the inner shaft 40 and the outer shaft 50provide a support and stiffness desired for operation of the gas turbineengine 20.

Referring to FIG. 7, another example shaft support configurationincludes the first and second forward bearing assemblies 70, 72 disposedto support the forward portion of the corresponding inner shaft 40 andouter shaft 50. The first aft bearing 74 is disposed aft of theconnection 80 between the rotor 78 and the inner shaft 40. The first aftbearing 74 is a roller bearing and supports the inner shaft 40 in astraddle configuration. The straddle configuration can requireadditional length of the inner shaft 40 and therefore an alternateconfiguration referred to as an overhung configuration can be utilized.In this example the outer shaft 50 is supported by the second aftbearing assembly 76 that is disposed forward of the connection 84between the high pressure turbine rotor 82 and the outer shaft 50.Accordingly, the connection hub 84 of the high pressure turbine rotor 82to the outer shaft 50 is overhung aft of the bearing assembly 76. Thispositioning of the second aft bearing 76 in an overhung orientationpotentially provides for a reduced length of the outer shaft 50.

Moreover the positioning of the aft bearing 76 may also eliminate theneed for other support structures such as the mid turbine frame 58 asboth the high pressure turbine 54 is supported at the bearing assembly76 and the low pressure turbine 46 is supported by the bearing assembly74. Optionally the mid turbine frame strut 58 can provide an optionalroller bearing 74A which can be added to reduce vibratory modes of theinner shaft 40.

Referring to FIGS. 8A and 8B, another example shaft supportconfiguration includes the first and second forward bearing assemblies70, 72 disposed to support corresponding forward portions of each of theinner shaft 40 and the outer shaft 50. The first aft bearing 74 providessupport of the outer shaft 40 at a location aft of the connection 80 ina straddle mount configuration. In this example, the aft portion of theouter shaft 50 is supported by a roller bearing assembly 86 supportedwithin a space 96 defined between an outer surface of the inner shaft 40and an inner surface of the outer shaft 50.

The roller bearing assembly 86 supports the aft portion of the outershaft 50 on the inner shaft 40. The use of the roller bearing assembly86 to support the outer shaft 50 eliminates the requirements for supportstructures that lead back to the static structure 36 through the midturbine frame 58. Moreover, the example bearing assembly 86 can provideboth a reduced shaft length, and support of the outer shaft 50 at aposition substantially in axial alignment with the connection hub 84 forthe high pressure turbine rotor 82 and the outer shaft 50. Asappreciated, the bearing assembly 86 is positioned aft of the hub 82 andis supported through the rearmost section of shaft 50. Referring to FIG.9, another example shaft support configuration includes the first andsecond forward bearing assemblies 70, 72 disposed to supportcorresponding forward portions of each of the inner shaft 40 and theouter shaft 50. The first aft bearing assembly 74 is supported at apoint along the inner shaft 40 forward of the connection 80 between thelow pressure turbine rotor 78 and the inner shaft 40.

Positioning of the first aft bearing 74 forward of the connection 80 canbe utilized to reduce the overall length of the engine 20. Moreover,positioning of the first aft bearing assembly 74 forward of theconnection 80 provides for support through the mid turbine frame 58 tothe static structure 36. Furthermore, in this example the second aftbearing assembly 76 is deployed in a straddle mount configuration aft ofthe connection 84 between the outer shaft 50 and the rotor 82.Accordingly, in this example, both the first and second aft bearingassemblies 74, 76 share a common support structure to the static outerstructure 36. As appreciated, such a common support feature provides fora less complex engine construction along with reducing the overalllength of the engine. Moreover, the reduction or required supportstructures will reduce overall weight to provide a further improvementin aircraft fuel burn efficiency.

Referring to FIG. 10, a portion of the example turbine section 28 isshown and includes the low pressure turbine 46 and the high pressureturbine 54 with the mid turbine frame 58 disposed between an outlet ofthe high pressure turbine and the low pressure turbine. The mid turbineframe 58 and vane 60 are positioned to be upstream of the first stage 98of the low pressure turbine 46. While a single vane 60 is illustrated,it should be understood these would be plural vanes 60 spacedcircumferentially. The vane 60 redirects the flow downstream of the highpressure turbine 54 as it approaches the first stage 98 of the lowpressure turbine 46. As can be appreciated, it is desirable to improveefficiency to have flow between the high pressure turbine 54 and the lowpressure turbine 46 redirected by the vane 60 such that the flow ofexpanding gases is aligned as desired when entering the low pressureturbine 46. Therefore vane 60 may be an actual airfoil with camber andturning, that aligns the airflow as desired into the low pressureturbine 46.

By incorporating a true air-turning vane 60 into the mid turbine frame58, rather than a streamlined strut and a stator vane row after thestrut, the overall length and volume of the combined turbine sections46, 54 is reduced because the vane 60 serves several functions includingstreamlining the mid turbine frame 58, protecting any static structureand any oil tubes servicing a bearing assembly from exposure to heat,and turning the flow entering the low pressure turbine 46 such that itenters the rotating airfoil 100 at a desired flow angle. Further, byincorporating these features together, the overall assembly andarrangement of the turbine section 28 is reduced in volume.

The above features achieve a more or less compact turbine section volumerelative to the prior art including both high and low pressure turbines54, 46. Moreover, in one example, the materials for forming the lowpressure turbine 46 can be improved to provide for a reduced volume.Such materials may include, for example, materials with increasedthermal and mechanical capabilities to accommodate potentially increasedstresses induced by operating the low pressure turbine 46 at theincreased speed. Furthermore, the elevated speeds and increasedoperating temperatures at the entrance to the low pressure turbine 46enables the low pressure turbine 46 to transfer a greater amount ofenergy, more efficiently to drive both a larger diameter fan 42 throughthe geared architecture 48 and an increase in compressor work performedby the low pressure compressor 44.

Alternatively, lower priced materials can be utilized in combinationwith cooling features that compensate for increased temperatures withinthe low pressure turbine 46. In three exemplary embodiments a firstrotating blade 100 of the low pressure turbine 46 can be a directionallysolidified casting blade, a single crystal casting blade or a hollow,internally cooled blade. The improved material and thermal properties ofthe example turbine blade material provide for operation at increasedtemperatures and speeds, that in turn provide increased efficiencies ateach stage that thereby provide for use of a reduced number of lowpressure turbine stages. The reduced number of low pressure turbinestages in turn provide for an overall turbine volume that is reduced,and that accommodates desired increases in low pressure turbine speed.

The reduced stages and reduced volume provide improve engine efficiencyand aircraft fuel burn because overall weight is less. In addition, asthere are fewer blade rows, there are: fewer leakage paths at the tipsof the blades; fewer leakage paths at the inner air seals of vanes; andreduced losses through the rotor stages.

The example disclosed compact turbine section includes a power density,which may be defined as thrust in pounds force (lbf) produced divided bythe volume of the entire turbine section 28. The volume of the turbinesection 28 may be defined by an inlet 102 of a first turbine vane 104 inthe high pressure turbine 54 to the exit 106 of the last rotatingairfoil 108 in the low pressure turbine 46, and may be expressed incubic inches. The static thrust at the engine's flat rated Sea LevelTakeoff condition divided by a turbine section volume is defined aspower density and a greater power density may be desirable for reducedengine weight. The sea level take-off flat-rated static thrust may bedefined in pounds-force (lbf), while the volume may be the volume fromthe annular inlet 102 of the first turbine vane 104 in the high pressureturbine 54 to the annular exit 106 of the downstream end of the lastairfoil 108 in the low pressure turbine 46. The maximum thrust may beSea Level Takeoff Thrust “SLTO thrust” which is commonly defined as theflat-rated static thrust produced by the turbofan at sea-level.

The volume V of the turbine section may be best understood from FIG. 10.As shown, the mid turbine frame 58 is disposed between the high pressureturbine 54, and the low pressure turbine 46. The volume V is illustratedby a dashed line, and extends from an inner periphery I to an outerperiphery O. The inner periphery is defined by the flow path of rotors,but also by an inner platform flow paths of vanes. The outer peripheryis defined by the stator vanes and outer air seal structures along theflowpath. The volume extends from a most upstream end of the vane 104,typically its leading edge, and to the most downstream edge of the lastrotating airfoil 108 in the low pressure turbine section 46. Typicallythis will be the trailing edge of the airfoil 108.

The power density in the disclosed gas turbine engine is much higherthan in the prior art. Eight exemplary engines are shown below whichincorporate turbine sections and overall engine drive systems andarchitectures as set forth in this application, and can be found inTable I as follows:

TABLE 1 Thrust Turbine section volume Thrust/turbine section Engine SLTO(lbf) from the Inlet volume (lbf/in³) 1 17,000 3,859 4.40 2 23,300 5,3304.37 3 29,500 6,745 4.37 4 33,000 6,745 4.84 5 96,500 31,086 3.10 696,500 62,172 1.55 7 96,500 46,629 2.07 8 37,098 6,745 5.50

Thus, in example embodiments, the power density would be greater than orequal to about 1.5 lbf/in³. More narrowly, the power density would begreater than or equal to about 2.0 lbf/in³. Even more narrowly, thepower density would be greater than or equal to about 3.0 lbf/in³. Morenarrowly, the power density is greater than or equal to about 4.0lbf/in³. Also, in embodiments, the power density is less than or equalto about 5.5 lbf/in³.

Engines made with the disclosed architecture, and including turbinesections as set forth in this application, and with modifications withinthe scope of this disclosure, thus provide very high efficientoperation, and increased fuel efficiency and lightweight relative totheir thrust capability.

An exit area 112 is defined at the exit location for the high pressureturbine 54 and an exit area 110 is defined at the outlet 106 of the lowpressure turbine 46. The gear reduction 48 (shown in FIG. 1) providesfor a range of different rotational speeds of the fan drive turbine,which in this example embodiment is the low pressure turbine 46, and thefan 42 (FIG. 1). Accordingly, the low pressure turbine 46, and therebythe low spool 30 including the low pressure compressor 44 may rotate ata very high speed. Low pressure turbine 46 and high pressure turbine 54operation may be evaluated looking at a performance quantity which isthe exit area for the respective turbine section multiplied by itsrespective speed squared. This performance quantity (“PQ”) is definedas:

PQ_(ltp)=(A _(lpt) ×V _(lpt) ²)  Equation 1:

PQ_(hpt)=(A _(hpt) ×V _(hpt) ²)  Equation 2:

where A_(lpt) is the area 110 of the low pressure turbine 46 at the exit106, V_(lpt) is the speed of the low pressure turbine section; A_(hpt)is the area of the high pressure turbine 54 at the exit 114, and whereV_(hpt) is the speed of the high pressure turbine 54.

Thus, a ratio of the performance quantity for the low pressure turbine46 compared to the performance quantify for the high pressure turbine 54is:

(A _(lpt) ×V _(lpt) ²)/(A _(hpt) ×V _(hpt)²)=PQ_(ltp)/PQ_(hpt)  Equation 3:

In one turbine embodiment made according to the above design, the areasof the low and high pressure turbines 46, 54 are 557.9 in² and 90.67in², respectively. Further, the speeds of the low and high pressureturbine 46, 54 are 10179 rpm and 24346 rpm, respectively. Thus, usingEquations 1 and 2 above, the performance quantities for the example lowand high pressure turbines 46,54 are:

PQ_(ltp)=(A _(lpt) ×V _(lpt) ²)=(557.9 in²)(10179 rpm)²=57805157673.9in² rpm²  Equation 1:

PQ_(hpt)=(A _(hpt) ×V _(hpt) ²)=(90.67 in²)(24346 rpm)²=53742622009.72in² rpm²  Equation 2:

and using Equation 3 above, the ratio for the low pressure turbinesection to the high pressure turbine section is:

Ratio=PQ_(ltp)/PQ_(hpt)=57805157673.9 in² rpm²/53742622009.72 in²rpm²=1.075

In another embodiment, the ratio is greater than about 0.5 and inanother embodiment the ratio is greater than about 0.8. WithPQ_(ltp)/PQ_(hpt) ratios in the 0.5 to 1.5 range, a very efficientoverall gas turbine engine is achieved. More narrowly, PQ_(ltp)/PQ_(hpt)ratios of above or equal to about 0.8 provides increased overall gasturbine efficiency. Even more narrowly, PQ_(ltp)/PQ_(hpt) ratios aboveor equal to 1.0 are even more efficient thermodynamically and from anenable a reduction in weight that improves aircraft fuel burnefficiency. As a result of these PQ_(ltp)/PQ_(hpt) ratios, inparticular, the turbine section 28 can be made much smaller than in theprior art, both in diameter and axial length. In addition, theefficiency of the overall engine is greatly increased.

Referring to FIG. 11, portions of the low pressure compressor 44 and thelow pressure turbine 46 of the low spool 30 are schematically shown andinclude rotors 116 of the low pressure turbine 46 and rotors 132 of thelow pressure compressor 44. The rotors for each of the low compressor 44and the low pressure turbine 46 rotate at an increased speed compared toprior art low spool configurations. Each of the rotors 116 includes abore radius 122, a live disk radius 124 and a bore width 126 in adirection parallel to the axis A. The rotors 116 support turbine blades118 that rotate relative to the turbine vanes 120. The low pressurecompressor 44 includes rotors 132 including a bore radius 134, a livedisk radius 136 and a bore width 138 in a direction parallel to the axisA. The rotors 132 support compressor blades 128 that rotate relative tovanes 130.

Referring to FIG. 12, with continued reference to FIG. 11, the boreradius 122 is that radius between an inner most surface of the bore andthe axis. The live disk radius 124 is the radial distance from the axisof rotation A and a portion of the rotor supporting airfoil blades. Thebore width 126 of the rotor in this example is the greatest width of therotor and is disposed at a radial distance spaced apart from the axis Adetermined to provide desired physical performance properties.

The increased speed of the low spool 30 as provided by the increasedspeeds of the disclosed compact turbine section 28 is provided by arelationship between the live disk radius 124 (r) to the bore radius 122(R) defined by a ratio of the live disk radius 124 over the bore radius122 (i.e., r/R). In the disclosed example embodiment the ratio isbetween about 2.00 and about 2.30. In another disclosed exampleembodiment, the ratio of r/R is between about 2.00 and 2.25.

The rotors 116 and 132 include the bore width 126 and 138 (W). The borewidths 126 and 138 are widths at the bore that are separate from a shaftsuch as the low shaft 40 of the low spool (FIG. 1). In one non-limitingdimensional embodiment, the widths 126, 138 (W) are between about 1.40and 2.00 inches (3.56 and 5.08 cm). In another non-limiting dimensionalembodiment, the widths 126, 138 (W) are between about 1.50 and 1.90inches (3.81 and 4.83 cm). Moreover, a relationship between the widths126, 138 (W) and the live rim radius 124 (r) is defined by ratio of r/W.In a disclosed example the ratio r/W is between about 4.65 and 5.55. Inanother disclosed embodiment the ratio of r/W is between about 4.75 andabout 5.50.

Accordingly, the increased performance attributes and performance areprovided by desirable combinations of the disclosed features of thevarious components of the described and disclosed gas turbine engineembodiments.

FIG. 13 shows an embodiment 200, wherein there is a fan drive turbine208 driving a shaft 206 to in turn drive a fan rotor 202. A gearreduction 204 may be positioned between the fan drive turbine 208 andthe fan rotor 202. This gear reduction 204 may be structured and operatelike the gear reduction disclosed above. A compressor rotor 210 isdriven by an intermediate pressure turbine 212, and a second stagecompressor rotor 214 is driven by a turbine rotor 216. A combustionsection 218 is positioned intermediate the compressor rotor 214 and theturbine section 216.

FIG. 14 shows yet another embodiment 300 wherein a fan rotor 302 and afirst stage compressor 304 rotate at a common speed. The gear reduction306 (which may be structured as disclosed above) is intermediate thecompressor rotor 304 and a shaft 308 which is driven by a fan driveturbine.

The embodiments 200, 300 of FIG. 13 or 14 may be utilized with thefeatures disclosed above.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of this disclosure. For that reason, the followingclaims should be studied to determine the scope and content of thisdisclosure.

1. A gas turbine engine comprising: a fan section including less than 18fan blades; a gear train; a compressor section; a combustor in fluidcommunication with the compressor section; a turbine section in fluidcommunication with the combustor and including a fan drive turbine and asecond turbine, the fan drive turbine driving a fan rotor through thegear train wherein: the fan drive turbine includes at least one rotorhaving a bore radius (R), and a live rim radius (r¹), and a ratio ofr¹/R is between 2.00 and 2.30, and at least one rotor having a borewidth (W) and a live rim radius (r²) and a ratio of r²/W is between 4.65and 5.55.
 2. The gas turbine engine set forth in claim 1, wherein the atleast one rotor with a ratio r¹/R between 2.00 and 2.30 is the samerotor as the at least one rotor with a ratio of r²/W between 4.65 and5.55.
 3. The gas turbine engine set forth in claim 1, wherein the secondturbine is a two stage turbine.
 4. (canceled)
 5. The gas turbine engineset forth in claim 3, wherein the gear train is a planetary gear system.6. The gas turbine engine set forth in claim 5, wherein the fan driveturbine has a first exit area and is configured to rotate at a firstspeed, the second turbine has a second exit area and is configured torotate at a second speed, which is faster than the first speed, saidfirst and second speeds being redline speeds, wherein a firstperformance quantity is defined as the product of the first speedsquared and the first area, a second performance quantity is defined asthe product of the second speed squared and the second area, and aperformance ratio of the first performance quantity to the secondperformance quantity is greater than or equal to 0.8 and less than orequal to 1.5.
 7. The gas turbine engine set forth in claim 6, whereinthe ratio of r¹/R is between 2.00 and 2.25.
 8. The gas turbine engine asset forth in claim 1, wherein the engine has a power density defined asSea Level Takeoff Thrust provided by the engine in lbf divided by avolume of the turbine section in in³ and the power density being greaterthan 1.5 lbf/in³ and less than 4.5 lbf/in³.
 9. (canceled)
 10. The gasturbine engine set forth in claim 1, wherein the gear train is aplanetary gear system, the fan drive turbine has a first exit area andis configured to rotate at a first speed, the second turbine has asecond exit area and is configured to rotate at a second speed, which isfaster than the first speed, said first and second speeds being redlinespeeds, and a first performance quantity is defined as the product ofthe first speed squared and the first area, a second performancequantity is defined as the product of the second speed squared and thesecond area, and a performance ratio of the first performance quantityto the second performance quantity is greater than or equal to 0.8 andless than or equal to 1.5.
 11. A gas turbine engine comprising: a fansection including a number of fan blades; a gear train; a compressorsection; a combustor in fluid communication with the compressor section;a turbine section in fluid communication with the combustor and theturbine section including a fan drive turbine driving a fan rotorthrough the gear train and a second turbine, wherein: the fan driveturbine includes at least one rotor having a bore radius (R), and a liverim radius (r¹), and a ratio of r¹/R is between 2.00 and 2.30, and atleast one rotor having a bore width (W) and a live rim radius (r²) and aratio of r²/W is between 4.65 and 5.55; the fan drive turbine has anumber of fan drive turbine stages, and a ratio between the number offan blades and the number of fan drive turbine stages is between 2.5 and8.5; and the fan drive turbine has a first exit area and is configuredto rotate at a first speed, the second turbine has a second exit areaand is configured to rotate at a second speed, which is faster than thefirst speed, said first and second speeds being redline speeds, and afirst performance quantity is defined as the product of the first speedsquared and the first area, a second performance quantity is defined asthe product of the second speed squared and the second area, and aperformance ratio of the first performance quantity to the secondperformance quantity is greater than or equal to 0.8 and less than orequal to 1.5.
 12. The gas turbine engine set forth in claim 11, whereinthe second turbine is a two stage turbine.
 13. (canceled)
 14. The gasturbine engine set forth in claim 12, wherein the number of fan bladesis less than 18 fan blades.
 15. The gas turbine engine set forth inclaim 14, wherein the at least one rotor with a ratio r¹/R between 2.00and 2.30 is the same rotor as the at least one rotor with a ratio ofr²/W between 4.65 and 5.55.
 16. The gas turbine engine set forth inclaim 14, wherein the engine has a power density defined as Sea LevelTakeoff Thrust provided by the engine in lbf divided by a volume of theturbine section in in³ and the power density being greater than 1.5lbf/in³ and less than 4.5 lbf/in³.
 17. (canceled)
 18. The gas turbineengine set forth in claim 11, wherein the at least one rotor with aratio r¹/R between 2.00 and 2.30 is the same rotor as the at least onerotor with a ratio of r²/W between 4.65 and 5.55.
 19. The gas turbineengine set forth in claim 11, wherein the number of fan blades is lessthan 18 fan blades. 20.-21. (canceled)
 22. A gas turbine enginecomprising: a fan section including a number of fan blades; a geartrain; a compressor section; a combustor in fluid communication with thecompressor section; a turbine section in fluid communication with thecombustor and including a fan drive turbine and a second turbine, thefan drive turbine driving a fan rotor through the gear train wherein:the fan drive turbine includes at least one rotor having a bore radius(R), and a live rim radius (r¹), and a ratio of r¹/R is between 2.00 and2.25, and at least one rotor having a bore width (W) and a live rimradius (r²) and a ratio of r²/W is between 4.65 and 5.55; and the fandrive turbine has a number of fan drive turbine stages, and a ratiobetween the number of fan blades and the number of fan drive turbinestages is between 2.5 and 8.5. 23.-24. (canceled)
 25. The gas turbineengine set forth in claim 22, wherein the number of fan blades is lessthan 18 fan blades.
 26. (canceled)
 27. The gas turbine engine set forthin claim 26, wherein the number of fan blades is less than 18 fanblades, the second turbine is a two stage turbine, and the gear train isa planetary gear system.
 28. The gas turbine engine set forth in claim22, wherein the number of fan blades is less than 18 fan blades, thesecond turbine is a two stage turbine, and the gear train is a planetarygear system.
 29. The gas turbine engine set forth in claim 22, whereinthe at least one rotor with a ratio r¹/R between 2.00 and 2.30 is thesame rotor as the at least one rotor with a ratio of r²/W between 4.65and 5.55.
 30. The gas turbine engine set forth in claim 22, wherein theratio of r²/W is between 4.75 and 5.50.
 31. The gas turbine engine asset forth in claim 8, wherein said power density being greater than orequal to 2.0 lbf/in³.
 32. The gas turbine engine as set forth in claim31, wherein the power density being greater than 3.0 lbf/in³.
 33. Thegas turbine engine as set forth in claim 32, wherein said power densitybeing greater than 4.0 lbf/in³.
 34. The gas turbine engine as set forthin claim 16, wherein said power density being greater than or equal to2.0 lbf/in³.
 35. The gas turbine engine as set forth in claim 34,wherein the power density being greater than 3.0 lbf/in³.
 36. The gasturbine engine as set forth in claim 35, wherein said power densitybeing greater than 4.0 lbf/in³.
 37. The gas turbine engine set forth inclaim 22, wherein the engine has a power density defined as Sea LevelTakeoff Thrust provided by the engine in lbf divided by a volume of theturbine section in in³ and the power density being greater than 1.5lbf/in³ and less than 4.5 lbf/in³.
 38. The gas turbine engine as setforth in claim 37, wherein the power density being greater than 3.0lbf/in³.
 39. The gas turbine engine as set forth in claim 38, whereinsaid power density being greater than 4.0 lbf/in³.